AOCS system to maintain planarity for space digital beam forming using carrier phase differential GPS, IMU and magnet torques on large space structures

ABSTRACT

A closed-loop motion monitoring and control system for structural mode control in a large, flexible space structure. The system uses combined sensor data to detect low-magnitude, low-frequency motion, estimate structure deformation constants, and damp structural vibrations with electromagnetic torque application.

RELATED APPLICATIONS

This application claims the benefit of priority of U.S. ProvisionalApplication No. 62/976,143, filed on Feb. 13, 2020, the content of whichis relied upon and incorporated herein by reference in its entirety.

FIELD

This disclosure relates to the fields of motion determination and motioncontrol for large, flexible space structures. Flexible structures arethose whose stiffness is low along one or more axes such that thestructure exhibits broad, slow differential displacements along thataxis when exposed to external forces and torques. The disclosureresolves challenges of measuring and controlling this motion. Thedisclosure includes a sensor suite for measuring displacements, a meansof integrating these measurements and estimating displacements in realtime, and control implementation.

BACKGROUND

U.S. Pat. No. 9,973,266 and U.S. Publ. No. 2019/0238216 show a systemfor assembling a large number of small satellite antenna assemblies inspace to form a large array. The entire content of the '266 patent isincorporated herein by reference. As disclosed in the '266 Patent, FIGS.1(a), 1(b) show a satellite communication system 100 having an array 300of common or small satellites 302 and a central or control satellite200. The small satellites 302 communicate with end users 500 within afootprint 400 on Earth, and also communicate with the control satellite200, which in turn communicates with a gateway 600 at a base station.The small satellites 302 can each include, for example, a processingdevice (e.g., a processor or controller) and one or more antennaelements. And the control satellite 200 can include a processing deviceand one or more antenna or antenna elements.

SUMMARY

A large array in space is formed by joining several smaller elements byconnectors, such as joints, hinges, tape-springs. Each element can beconsidered as rigid, flexure being largely in the connectors thatconnect the elements to each other. The connectors have a storageconfiguration in which they are bent so that the antenna elements 300are folded upon each other to be compact for transport into space. Andthe connectors have default bias to a deployed configuration in whichthey expand so that the antenna elements are unfolded and expand into alarge planar configuration in space. In addition, the control satellite200 need not be distinct from the small satellites 302, but rather thecontrol satellite 200 can be connected to the small satellites 302, suchas directly embedded within the array 100.

However, low mass-per-unit aperture arrays can physically bend ordeviate from their nominal positions due to external forces in deployedorbit around the Earth (e.g., low earth orbit (LEO), medium earth orbit(MEO), etc.). For example, there can be both displacement and rotationin the connectors. Analysis has revealed that there could be as much as70 cm displacement in an 8 m diameter array, which can be corrected todisplacement of 10 cm or less by appropriate mechanical or structuralcompensations. This can be achieved by use of, e.g., torque rods thatapply a magnetic moment against the Earth's magnetic field which movesthe connectors toward their fully deployed configuration and moves thearray of antenna elements toward the full planar configuration. Residualdisplacement (after the mechanical compensation) is compensated bybeamforming corrections. Thus, the structural compensation describedherein applies a coarse correction, whereas the phase adjustment appliesa fine correction.

Accordingly, a system and method are provided to monitor and correct formotions caused by various mechanical modes of deviation to reducestructural deflections that may affect beamforming.

BRIEF DESCRIPTION OF THE FIGURES

FIGS. 1(a), 1(b) show a known phased array.

FIG. 2 is a block diagram of a structure.

FIG. 3(a) is a perspective view of a large phased array formed byintegrating several small satellites in space.

FIG. 3(b) shows a front view, top view, and right-side view of the arrayof FIG. 3(a).

FIG. 4(a) is a perspective view showing bending of a satellite antennaarray due to external forces.

FIG. 4(b) shows a front view, top view and right-side view of the arrayof FIG. 2(a).

FIG. 5(a) is a perspective view showing bending of a satellite antennaarray due to external forces.

FIG. 5(b) shows a front view, top view and right-side view of the arrayof FIG. 3(a).

FIG. 6 shows radiation patterns.

FIGS. 7(a)-7(c) show in-plane and out-plane tilts that may be present ineach small satellite.

FIGS. 8(a), 8(b) show relative in-plane and out-plane displacements thatmay be present across small satellites.

FIG. 9 is a block diagram overview of one embodiment of the presentdisclosure.

FIG. 10 is a block diagram of the estimator of FIG. 2 .

FIG. 11 is a perspective view of the system.

DETAILED DESCRIPTION

In describing the illustrative, non-limiting embodiments of thedisclosure illustrated in the drawings, specific terminology will beresorted to for the sake of clarity. However, the disclosure is notintended to be limited to the specific terms so selected, and it is tobe understood that each specific term includes all technical equivalentsthat operate in similar manner to accomplish a similar purpose. Severalembodiments of the disclosure are described for illustrative purposes,it being understood that the disclosure may be embodied in other formsnot specifically shown in the drawings.

Referring to FIG. 2 , in one example embodiment, only the electronics ofa single small satellite 302 is shown. In this embodiment the structure10 is an assembly or common small satellite that is connected to otherantenna assemblies in a large antenna array 5 (FIG. 3 ), such as in theantenna assembly 300 and array 100 of the '266 Patent (FIG. 1(a)). Theoverall system forms an Altitude and Orbit Control System (AOCS) thatcan include the common satellite 10, control satellite 200, and/orground station. The control satellite 200 can be fixedly connected tothe small satellites 302, such as at the center of the array as in 5(FIG. 3(a)).

The structure 10 is flat and rectangular or square, with thecommunication components (e.g., antenna elements 19) at one side surface7 (FIG. 3(a)) facing the Earth (nadir) to communicate with user devices(e.g., cell phones) and an opposite side surface 9 (FIG. 3(a)) facing inthe opposition direction (zenith) with solar cells 17 that generatesolar power for use by the electronic components, e.g., a processingdevice 12, antennas 19, battery 16, and antenna front end modules 18.

Each structure 10 also has one or more connectors 14, such as a hinge,joint, spring or tape-spring connector, that connect the structure 10 toone or more neighboring similar structures 10. As shown, the structure10 can be rectangular or square and encompass multiple antenna elements300, and one or more connectors can be positioned at or along one ormore of the edges or sides of the structure 10. It will be recognizedthat the system can utilize any suitable connection system, for examplesuch as the one shown and described in U.S. Patent Pub. Nos.2020/0361635, 2020/0366237, and 2020/0365966, the entire contents ofwhich are hereby incorporated by reference.

The connectors can be subject to bending or flexing in the operatingconfiguration. For example, the maximum flex at the connector 14 mightbe several degrees. Any flex results in a deviation of the antennaelements from the planar configuration in which the communication side(and/or the solar side) of the plurality of antenna elements are planar.That deviation is undesirable since it can affect beam formation.

In this example embodiment, it is therefore desirable for the array 5 tobe substantially flat on both the solar side and the communication side,i.e., that the individual structures 10 are flat on both sides and thatthey are planar or co-planar with one another on both sides so that theoverall array 5 is planar on both sides. However, the structure 10and/or array 5 is subject to forces in space that can cause thestructure 10 or array 5 to flex or bend.

FIGS. 3(a), (b) show a nearly circular large planar phased array 5formed in space by integrating many small satellite structures 10. Eachsmall satellite could host a processing device 12 (e.g., processor) andseveral antenna elements 19. Hundreds of such small satellites togethercould form a large phased array with thousands of overall antennaelements. In one embodiment, each small satellite (referred to here as amicron) has an antenna assembly with antenna elements 19 arranged infour rows and four columns in a square shape. The overall phased arrayformed by the interconnection of several small satellites could take asquare, or a rectangular or a circularized shape as desired by theapplication. Any suitable small satellite can be utilized, such as shownand describe in U.S. Pub. Nos. 2020-0361635, 2020-0366237, and2020-0365966, the entire contents of which are hereby incorporated byreference.

The antenna elements 19 are positioned at the communication side surface7 of the array and the solar cells are positioned at the solar side 9 ofthe array 5. The arrow 11 shows the boresight, which for a planar phasedarray refers to a normal to the array's plane. Any beam off-boresight iscalled an edge beam (e.g., FIG. 6 )

The small satellites 10 communicate with end user devices (such as cellphones) on Earth, and with the central processor 200, which in turncommunicates with a gateway at the ground station. The signalscommunicated to/by the small satellites are aggregated together, suchthat the small satellites collectively transmit and receive signals tothe end user devices. However, any bending or flexing of the array cancause the signals from the individual small satellites to deviate or beout of phase from the desired phase.

FIGS. 4(a), 4(b), 5(a), 5(b) show the gradual bending effect in largearrays from the center towards the edge of the arrays. The individualantenna assemblies are rigid, but are mechanically coupled to oneanother by the connectors 14. Those connectors hold the small satellitestogether, but are subject to bending or flexing, and tend to oscillateinwards and outwards at low frequency, with maximum displacements at theextremities, depending on the external forces. As the mass per unitaperture is reduced, the stiffness of the array is reduced and the arrayencounters greater flexure. The arrows in FIGS. 4(a), 5(a) point to theboresight of the array.

FIGS. 4(a), 4(b) illustrate the inward (towards the arrow 11 orboresight) flexing of the array, and FIGS. 5(a), 5(b) illustrate theoutward (away from the arrow 11 or boresight) flexing of the array. Thearray is imagined to be in a nominal plane that is normal orperpendicular to the boresight. The flexing causes deviation from thenominal plane.

The example of FIGS. 4, 5 , shows a single bend 20 in the array 5.However, other bends are possible, for example bends that only partiallyextend along the array, bends that are offset from the center diameter,bends that extend at other positions and locations that do not passthrough the center of the array, and multiple bends. And, while the bend20 is shown having a sharp angle between the left and right halves 22,24, the bend can be more curved. And, the left and right halves 22, 24need not be planar, but can be curved due to slight deviations or bendsat connectors between antenna modules 10.

Referring to FIG. 6 , the top row of measurements shows the expectedradiation patterns from a 10.3 m planar array of FIG. 3 while formingthe beams towards the boresight of the array and edge of the footprint.The bottom row of measurements in FIG. 6 shows the expected radiationpatterns for an array of FIGS. 4, 5 with 8.7° bending from nominal planewithout any compensation for change in antenna elements' position whilebeamforming. As shown, there is a distorted radiation pattern with dualmain-lobe and the reduced array gain due to bending of the array.

While FIGS. 4, 5 show uniform bending effect, there could be randomperturbations in the position of each small satellite structure 10 whiledeploying and attaching to neighboring small satellites. FIG. 7(a) showsa nominal orientation for a structure 10, and FIGS. 7(b)-7(c), 8(a)-8(b)show such perturbations in the form of in-plane and out-plane tilts anddisplacements for a plurality of antenna elements. FIG. 7(a) and the topdrawings of FIGS. 7(b), 7(c), each show a single structure 10 (which mayhave multiple antenna elements). FIGS. 8(a), 8(b), and the bottomdrawings in FIGS. 7(b), 7(c) show multiple structures 10 coupled to oneanother by connectors, and bending at each connector. Generaldisplacements such as those in FIG. 8 are not controllable as describedin the current embodiment, but such displacements are observable usingthe determination methods described herein if they are sufficientlylarge. Thus, FIGS. 4, 5, 7, 8 depict the main types of perturbationsthat are likely to occur, any of which may increase the distortion ofthe radiation pattern. The embodiment described herein has significantutility for near-planar arrays but may be utilized in systems withlarger non-planarity if a majority of such non-planarity is the resultof modal perturbations.

FIG. 9 is one embodiment of the disclosure in which the structure 10 inFIG. 7 has a single Global Positioning System (GPS) unit that reportscarrier phase data. A collection of such antenna elements collectivelycontain a set of GPS units 50 whose data is used to calculated a CarrierPhase Differential GPS (CD-GPS) array 51 of differential GPS solution212. Each structure 10 also includes an Inertial Measurement Unit 49,and the collective IMU data is compiled in an array 52 of filtered IMUmeasurements 515. Such IMUs may be inertial sensing units whichintegrate acceleration only and/or directly detect angular motion, orother such inertial measurement devices.

The embodiment in FIG. 9 also includes a structure estimator 54 acomputed torque model 62 based on that structure estimator, a torquecommand 66 issued to a torque mechanism, and a feedback element 64 fromthat torque mechanism. In one embodiment, the GPS sensors 50 arepositioned as close as possible to the IMU sensors 49. The GPS sensors50 and/or IMU sensors 49 are coupled to the structure elements 10; eachstructure 10 can have one or more GPS sensor 50 and one or more IMUsensor 49, or one GPS sensor 50 and/or one IMU sensor 49 can beassociated with multiple structures 10. In one embodiment, the IMUsensors 49 are accelerometers that detect an acceleration.

The effect of modal array deformation caused by coupled flexure onbeamforming can be minimized by considering the instantaneous positionof each antenna element while computing the corresponding phases usedfor beamforming. This is accomplished by placing position sensors 50(FIGS. 2, 7 (a)) and IMUs 49 in several of the small satellites 10 foraccurate position and rate estimation of each antenna element or minimumrequired sensors, to predict the uniform perturbation characteristic ofthe small satellites across array, to determine approximate position ofeach small satellite. For example, each small satellite 10 can have oneor more sensors that are placed on the body of the structure, such as onthe communication side (the side with the antennas 19) or the solar side(the side with the solar cells 17). The sensor placement is determinedby the type of sensor to ensure observability of the sensed effect. Inthe current embodiment, the GPS antenna attached to the receiver 50 mustface the GPS constellation and therefore must be on the solar side ofthe antenna element structure 10, while the IMU may be placed near thecenter of the structure 10.

As shown in FIG. 11 , each pair of GPS receivers must observeoverlapping signals from the GPS constellation. The sensors can, forexample, be part of the electronic circuits that form the smallsatellite. Reference X, Y and Z planes for a nominal array are shown inFIG. 7(a). The sensors can be a standard Global Positioning System (GPS)receivers or other sensor devices that automatically estimate positionin a global co-ordinate system to the accuracy of carrier phasedifferential GPS (approx. 2 cm relative accuracy).

The sensor 50 in this embodiment is a standard Global Positioning System(GPS) sensor device that automatically estimates position in a globalco-ordinate system and provides carrier phase and pseudorange outputdata for individual received GPS signals. As shown in FIG. 11 , the GPSreceivers 50 reside at different locations 298 and 299 on the satellite202. Each receiver senses its position 160 relative to common GPSsatellites (exemplified here as 135, 137), then use carrier phaseinformation 152 as it changes 154 over time 150 to determine relativelocation 170 of the receivers. A group of such sensors 50 with mutuallycommon received signals results in a set of solutions 51 providingrelative offsets 170 between the receivers 50. This process is known asCD-GPS and is well-documented in academic and technical papers, thoughany suitable technique can be utilized. In order to perform the systemoperations described herein, the number of CD-GPS solutions 51 need onlyexceed some minimum requirement. There is no theoretical maximum to thenumber that may be used, as additional solutions tend to improve thesolution.

The embodiment here further utilizes inertial measurement unit (IMU)sensors 49 that provide measurements of the motion of a subset ofsatellite elements 52. The raw IMU data is filtered 84 (FIG. 9 ) withe.g. a low-pass filter to produce a dataset with content at a higherrate than the structure estimator operation and with low-quality oroutlying datapoints removed. Such IMU data is not required to coincideidentically in time with the available CD-GPS data, but the sensors mustbe placed to capture oscillatory motion within the system. As with theGPS solutions, the number of IMU datapoints need only exceed someminimum requirement for system observability, but no theoretical maximumnumber exists.

Timing misalignments between CD-GPS and IMU solutions are resolved viapropagation of the existing solution. These CD-GPS and IMU solutions arethe inputs to a structure filter 83 that estimates the characteristicparameters describing the displacements of the spacecraft structuralelements, as well as the persistent error (bias estimate) in the IMUrate measurements 85. These bias estimates are applied to the availablerate data solution 60 even if new CD-GPS data is not available. Duringthese periods, the structure motion is propagated 87 solely using IMUdata to provide an estimate of the current location of each spacecraftelement 90. Structure constants may be updated at each timestep 88 toreflect the most recent observations, or if no control is performed theymay be updated only when CD-GPS is available. Note that the solutionalso requires as an input the baseline angular motion of the spacecraftsystem 89. Once structural parameters are estimated, the positions X, Y,and Z of the structure elements can be determined as often as requiredto provide data to beamforming. Instantaneous solution accuracy issub-centimeter and can be extrapolated to similar accuracy over veryshort timespans.

The placements of the sensors 49 and 50 in the example realization arechosen to coincide with the maximum displacements due to the primaryoscillation modes of the aperture structure. Because the beamformingphase is used to determine the phase compensation based on thedisplacement of the structure elements 10 from a planar configuration,the system must compensate for any additional flexing by mechanicallycorrecting the relative positioning of the spacecraft elements 302 wherepossible. Where each small satellite 10 has its own sensors 49 and 50,the deviation can be determined based on the position of that sensor.However, where a single sensor is provided for multiple smallsatellites, the deviation and correction for each small satellite 10 canbe interpolated based on the positions of surrounding small satelliteswith respect to the sensor. In the embodiment shown here, thesedisplacements are determined at the central satellite 200 (FIG. 1(a)) bycalculating the apparent modal excitation 88. Determining thedisplacements resulting from these modes offers information that can beused to counteract their effects through a variety of means, anddirectly controlling these modes can improve spacecraft performance.

The present system resolves that measurement problem with a real-timeestimation algorithm and integrates the solution with space-capableactuators to perform closed-loop control. The displacement filter 56outputs the displacement data to the correction and control module 58,which uses an actuator 23 operating on an external field. In oneembodiment, the actuator 23 can be an electromagnetic torque rod thatapplies a physical displacement to the structure 10 and correct theoverall displacement of the array 5, reducing deflections about thebending axis 20. The correction and control module 58 uses current andextrapolated deformation constant estimates and determines apparentexcitation energy. It then applies the correction to the actuator 23 bya damping torque model 62, which sends a torque command 66 to the torquerod.

The torque rod then applies a torque to the small satellite 10 to moveit back toward the desired position at the point of maximum deflection,and the resultant torque is fed back to the model 64. In this way thepeak displacement decays. The control law associated with the estimatedCD-GPS/IMU mode dynamics can take any basic nonlinear form, or a linearform if the structure is sufficiently rigid. Because mode motion isoscillatory as a function of the estimated mode shape constants 88, andbecause digital compensation for some limited motion is possible, theselected control law need only force the system to remain within anallowable equilibrium displacements.

Thus, the processing device at the central controller 200 computes thecommanded actuator input to move one or more of the plurality of antennaelement structures 10 to correct for structural displacement of theantenna element structures. This computation may include prior knownactuator command 64, which can affect system dynamics and requiredtorque inputs going forward. The correction and control module 58 alsooutputs the displacement data (e.g., the sub-centimeter displacement) tothe digital beamformer 68. The digital beamformer 68 computes andapplies a phase correction to the beam. Any suitable technique can beutilized to apply an electronic correction, such as described inco-pending application no., which claims priority to 62/976,107, theentire contents of which are hereby incorporated by reference.

As a more specific example of the embodiment described herein, forsystems with a significant first bending mode (see FIGS. 3-4 ), bendingof the longest beam of the system about a perpendicular eigenaxis can alocally linearized as a 1-dimensional Euler-Bernoulli beam. Bending isapproximately quadratic. The controller is designed only to stabilizethe principal modes, so a constraint is placed on the energy imparted tothe system (|u|<|u_(max)|) to prevent possible excitation of additionalmodes. As noted above, control is imparted symmetrically about theinertial eigenaxes. Deformation energy is driven to 0, with a hysteresisenacted to limit chatter as the allowed maximum deformation is achieved.

Thus, as described above, the present system determines the amount oflocal movement of the structural array 30, and then corrects that. It isfurther noted that co-pending application no., which claims priority to62/976,107, determines the amount of flexing or bending of thestructural array 5, and then corrects that by performing beam formingtechniques that compensate for the bending. The entire content of thatapplication is incorporated herewith.

In one example embodiment, structure 10 is an antenna assembly with asolar panel that receives solar energy from the Sun and generates solarpower for use by the structure. The overall structure is flat andrectangular or square, such as a tile, with the communication components(e.g., antenna elements) at one side surface facing the Earth (nadir) tocommunicate with user devices (e.g., cell phones) and an opposite sidesurface facing in the opposition direction (zenith) with solar cellsthat generate solar power for use by the electronic components—e.g., aprocessing device, antennas, antenna front end modules. Here the controlsatellite 200 is fixedly connected to the small satellites 302, shown atthe center of the array 100 and visible in the array 5.

It is important that the structure 10 and the array 5 remain as flat(i.e., planar) as possible to maximize solar power generation by thesolar side and communication with the Earth on the communication side.Thus, it is desirable for the array 5 to be substantially flat, i.e.,that the structures 10 are flat and that they are planar with oneanother. However, the structure 10 and/or array 5 is subject to forcesin space that can cause the structure 10 or array 5 to flex or bend. Tocorrect for any bending or flexing, the structure 10 hassymmetrically-placed GPS units 50, inertial measurement units 49, andactuators 23 on each small spacecraft.

The processing device 12 in the common satellite 10 transmits requireddata to the control satellite 200 in real time, which derives theinitial estimates of small displacements due to modes of the structureflex 11 via the CD-GPS solution 51 and IMU data 52 at time t=0. Thecontrol satellite 200 then determines the apparent effect of the bendingof the structure 11 and estimates the modal contributions of theexpected principal modes via structure constants. Approximately 1/10second later, additional filtered IMU data 52 is received, andcalculated bias and the spacecraft angular rate 89 are removed 86 fromthe sensed rate. The system is propagated forward 87. Structureconstants may be updated 88 if acceleration or rate data suggests alarge error in the estimate, resulting in a final displacement estimate90 and 58.

This data is used as an input to the damping torque computation 62which, along with any previous output torque 64, results in a newdesired torque command 66. These torque commands are sent to theappropriate small satellite actuators 23 as torque rod activationsignals. Finally, the estimate 90 and 58 is sent to the beamformer 68,which uses the data to improve its digital beamforming solution. Thisprocess repeats for 0.25 seconds until new CD-GPS data is available, atwhich point the estimated structure constants are updated using theposition data.

When the structure 10 is configured as an antenna array 5, it (e.g.,antenna 19 or antenna elements 302) communicates with processing deviceson Earth, such as for example a wireless device including a user device(e.g., cell phone, tablet, computer) and/or a ground station. Thepresent disclosure also includes the method of utilizing the structure10 to communicate with processing devices on Earth (i.e., transmitand/or receive signals to and/or from). The present disclosure alsoincludes the method of processing devices on Earth communicating withthe structure 10 (i.e., transmit and/or receive signals to and/or from).In addition, while the structure 10 is used in Low Earth Orbit (LEO) inthe examples disclosed, it can be utilized in other orbits or for otherapplications.

Still further, while the system has been described as for an array ofantenna assemblies, the system can be utilized for other applications,such as for example data centers, telescopes, reflectors, and otherstructures, both implemented in space or terrestrially. The system ofthe present disclosure can also be utilized in combination with a phasecorrection system, such as shown and described in U.S. application Ser.No. 17/175,428, filed herewith, entitled Compensating Oscillations in aLarge-Aperture Phased Array Antenna, claiming priority to U.S.Application No. 62/976,107, filed Feb. 13, 2020, the entire contents ofwhich are hereby incorporated by reference.

In addition, it is noted that operation is described as occurring at thecontrol satellite 200, which may or may not be fixedly embedded in thearray. However, operation can also be at the common satellite 10processing device 12 if GPS and IMU data from other structures 10 isdistributed in such fashion. In another embodiment of the presentdisclosure, data (such as position and attitude) can be transmitted fromthe satellite 10 and/or 200 (e.g., by the common satellite processingdevice 12 and/or the control satellite processing device, if such arenot coincident) to a ground station. The ground station processingdevice can then determine the necessary correction and/or other flightinformation and transmit a control signal to the satellite 10 and/or 200(e.g., common satellite processing devices 12 and/or control satelliteprocessing device) to control the correction via the torque rod 23, inaddition to performing other ground-based tasks.

It is further noted that the drawings may illustrate and the descriptionand claims may use several geometric or relational terms and directionalor positioning terms, such as planar, linear, curved, circular, flat,left, and right. Those terms are merely for convenience to facilitatethe description based on the embodiments shown in the figures, and arenot intended to limit the disclosure. Thus, it should be recognized thatthe system can be described in other ways without those geometric,relational, directional or positioning terms. In addition, the geometricor relational terms may not be exact. For instance, walls or surfacesmay not be exactly flat, or planar to one another but still beconsidered to be substantially planar because of, for example, roughnessof surfaces, tolerances allowed in manufacturing, etc. And, othersuitable geometries and relationships can be provided without departingfrom the spirit and scope of the disclosure.

The foregoing description and drawings should be considered asillustrative only of the principles of the disclosure. The system may beconfigured in a variety of shapes and sizes and is not intended to belimited by the embodiment. Numerous applications of the system willreadily occur to those skilled in the art. Therefore, it is not desiredto limit the disclosure to the specific examples disclosed or the exactconstruction and operation shown and described. Rather, all suitablemodifications and equivalents may be resorted to, falling within thescope of the disclosure.

The invention claimed is:
 1. A system comprising: a plurality ofstructures each having a planar surface, the plurality of structuresconnected to form a structural array with a planar array surface; asensor coupled to one or more structures of the plurality of structuresand configured to detect a position of the one or more structures of theplurality of structures; a processing device configured to monitorstructural displacement modes of the plurality of structures or thestructural array and determine a correction; and an actuator configuredto receive the correction from the processing device and apply thecorrection to the plurality of structures or the structural array,wherein the actuator is a magnetic torque rod configured to generate alocal magnetic field against an Earth magnetic field and to apply torqueto the plurality of structures or the structural array, and wherein datacorresponding to the applied torque is to be transmitted back to theprocessing device.
 2. The system of claim 1, wherein the plurality ofstructures each comprises an antenna assembly.
 3. The system of claim 1,wherein the plurality of structures are flat, and square or rectangular.4. The system of claim 1, wherein the plurality of structures comprisesan antenna assembly, and wherein the structural array comprises anantenna array.
 5. The system of claim 1, wherein the plurality ofstructures is configured to communicate with a processing device onEarth.
 6. The system of claim 1, wherein the sensor comprises at leastone of a Global Positioning System (GPS) receiver or an InertialMeasurement Unit (IMU).
 7. The system of claim 1, wherein the processingdevice is configured to monitor structural displacement modes bydetermining a position and velocity difference between GlobalPositioning System (GPS) receivers based on carrier-phase differentialGPS (CD-GPS) measurements.
 8. A system comprising: a plurality ofstructures each having a planar surface, the plurality of structuresconnected to form a structural array with a planar array surface; aprocessing device configured to monitor structural displacement modes ofthe plurality of structures or the structural array and determine acorrection; and an actuator configured to receive the correction fromthe processing device and apply the correction to the plurality ofstructures or the structural array, wherein the actuator is a magnetictorque rod configured to generate a local magnetic field against anEarth magnetic field and apply torque to the plurality of structures orthe structural array, wherein data corresponding to the applied torqueis to be transmitted back to the processing device, and wherein theprocessing device is configured to control planarity of the plurality ofstructures or the structural array in real time.
 9. The system of claim8, further comprising a sensor coupled to one or more of the pluralityof structures and configured to detect a position of the one or morestructures of the plurality of structures.
 10. The system of claim 9,wherein the sensor comprises at least one of a Global Positioning System(GPS) receiver or an Inertial Measurement Unit (IMU).
 11. The system ofclaim 8, wherein the plurality of structures are flat, and square orrectangular.
 12. The system of claim 8, wherein the plurality ofstructures comprises an antenna assembly, and wherein the structuralarray comprises an antenna array.
 13. The system of claim 8, wherein theplurality of structures is configured to communicate with a processingdevice on Earth.
 14. The system of claim 8, wherein the processingdevice is configured to monitor structural displacement modes bydetermining a position and velocity difference between GlobalPositioning System (GPS) receivers based on carrier-phase differentialGPS (CD-GPS) measurements.
 15. A system comprising: a plurality ofstructures each having a planar surface, the plurality of structuresconnected to form a structural array with a planar array surface; asensor coupled to one or more of the plurality of structures andconfigured to detect a position of the one or more structures of theplurality of structures, the sensor comprising one or more GlobalPositioning System (GPS) receivers and Inertial Measurement Units(IMUs); a processing device configured to monitor structuraldisplacement modes of the plurality of structures or the structuralarray, and determine a correction, wherein the processing device isconfigured to monitor structural displacement modes by determining aposition and velocity difference between GPS receivers based oncarrier-phase differential GPS (CD-GPS) measurements; and an actuatorconfigured to receive the correction from the processing device andapply the correction to the plurality of structures or the structuralarray, wherein the actuator is a rod configured to apply torque to theplurality of structures or the structural array, and wherein datacorresponding to the applied torque is to be transmitted back to theprocessing device.
 16. The system of claim 15, wherein the actuatorcomprises a magnetic torque rod configured to generate a local magneticfield against an Earth magnetic field.
 17. The system of claim 15,wherein the plurality of structures are flat, and square or rectangular.18. The system of claim 15, wherein the plurality of structurescomprises an antenna assembly, and wherein the structural arraycomprises an antenna array.
 19. The system of claim 15, wherein theplurality of structures is configured to communicate with a processingdevice on Earth.
 20. The system of claim 15, wherein the processingdevice is configured to control planarity of the plurality of structuresor the structural array in real time.